Turbine engine assembly with a component having a leading edge trough

ABSTRACT

An assembly for a turbine engine comprises a circumferential band defining an air flow path, a component having a leading edge and extending from the band into the flow path, and a trough located in a surface of the band along the leading edge.

BACKGROUND OF THE INVENTION

Turbine engines, and particularly gas or combustion turbine engines, arerotary engines that extract energy from a flow of combusted gasespassing into the engine through a series of compressor stages, acombustor, and a series of turbine stages where each compressor stageand turbine stage includes a multitude of rotating blades and stationaryvanes. Turbine engines are commonly used for aeronautical applicationssuch as propulsion of aircraft, and also for terrestrial applicationssuch as power generation.

Turbine engines for aircraft utilize mainstream flow to drive therotating turbine blades to generate thrust. The mainstream flow ispropelled by combustion of gas to increase the thrust generated by theengine, the mainstream flow can create a bow wave in front of theturbine airfoils.

Turbine engines for aircraft are designed to operate at hightemperatures to maximize engine efficiency. Temperatures in the highpressure turbine can be around 1000° -2000° C., and the spacing betweeneach vane and blade in a stage can be constructed to prevent ingestionof the bow wave into regions that are sensitive to high temperatures.

BRIEF DESCRIPTION OF THE INVENTION

In one aspect, an assembly for a turbine engine comprises radiallyspaced inner and outer bands that define a flow path between the bands,where the inner and outer bands each have an upstream edge, a downstreamedge, and a surface extending between the upstream and downstream edges.The assembly further comprises a component extending from at least oneof the inner and outer bands into the flow path and defining a leadingedge confronting the flow path, and also comprises a trough located inthe surface of the at least one of the inner and outer bands along theleading edge.

In another aspect, a turbine assembly for a turbine engine comprises anouter band, an inner band radially spaced inwardly from the outer bandand at least one of the outer and inner bands having an upstream edgeand a downstream edge with a surface extending between, a vane having anouter wall extending from a leading edge to a trailing edge to define achord-wise direction and extending from a root to a tip to define aspan-wise direction, with the root proximate the surface of theplatform, and a trough located in the upper surface at least along theleading edge.

In yet another aspect, a method of controlling the upstream extent of abow wave from an airfoil in a gas turbine engine comprises forming avortex at a leading edge of the airfoil along at least a portion of theroot of the airfoil.

BRIEF DESCRIPTION OF THE DRAWINGS

In the drawings:

FIG. 1 is a schematic cross-sectional diagram of a gas turbine enginefor an aircraft.

FIG. 2 is a side view of a portion of a high pressure turbine assemblyof the turbine engine in FIG. 1, with the stage comprising rotatingairfoils in the form of blades and stationary airfoils in the form ofvanes.

FIG. 3 is a closeup side view of the turbine assembly of FIG. 2.

FIG. 4 is a perspective view of one of the airfoils, which has beensimplified into an aerodynamic wedge of the turbine assembly in FIG. 2according to a first embodiment of the invention.

FIG. 5 is a side sectional view of a portion of the airfoil in FIG. 4taken along line 5-5.

FIG. 6 is a top view of the airfoil in FIG. 4.

FIG. 7 is a perspective view of one of the airfoils of the turbineassembly in FIG. 2 according to a second embodiment of the invention.

FIG. 8 is a side view of a portion of the airfoil in FIG. 7 taken alongline 8-8.

DESCRIPTION OF EMBODIMENTS OF THE INVENTION

A gap or cavity between flow path components in a turbine enginetypically contains materials that are sensitive to high temperatures,and it is beneficial to purge such cavities with cooler air. The cavitypurge pressure is set by the inner or outer band static pressure in thegas flow path. These gaps are subject to pressure variations in the flowpath, such as a bow wave that emanates from the lead edge of flow pathobstructions such as airfoils. The bow wave generates a locally highpressure which can result in ingestion of hot gases into the cavity thatcontains temperature-sensitive materials.

The bow wave strength and broadcast is driven by the flow path approachvelocity and airfoil lead-edge diameter. The approach velocity andlead-edge diameter are typically designed for optimal aerodynamicperformance, and therefore other methods are often evaluated to reducebow wave broadcast for a given aerodynamic design.

This invention seeks to reduce the forward broadcast of a bow wave froman airfoil lead edge or other flow path obstruction by the placement ofa trough around the leading edge of the airfoil at the interface witheither the inner or outer band. The trough can create a controlledvortex within the confines of the trough and suppress the bow wave inthe vicinity of the trough. The suppressed bow-wave can reduce oreliminates the ingestion of hot gases into the cavities between flowpath components and therefore allow the flow path axial lengths to bereduced, resulting in weight savings and lower frictional losses.

The described embodiments of the present invention are directed to atrough for an airfoil located along the leading edge of the airfoil in aturbine engine. For purposes of illustration, the present invention willbe described with respect to a vane in the turbine section of anaircraft turbine engine. It will be understood, however, that theinvention is not so limited and may be applied to a vane or bladelocated within a compressor section as well as the turbine section, andin the case of vanes can be applied to the inner or outer bands.Further, the present invention may have general applicability within anengine as well as in non-aircraft applications, such as other mobileapplications and non-mobile industrial, commercial, and residentialapplications.

As used herein, the term “forward” or “upstream” refers to moving in adirection toward the engine inlet, or a component being relativelycloser to the engine inlet as compared to another component. The term“aft” or “downstream” used in conjunction with “forward” or “upstream”refers to a direction toward the rear or outlet of the engine or beingrelatively closer to the engine outlet as compared to another component.

Additionally, as used herein, the terms “radial” or “radially” refer toa dimension extending between a center longitudinal axis of the engineand an outer engine circumference.

All directional references (e.g., radial, axial, proximal, distal,upper, lower, upward, downward, left, right, lateral, front, back, top,bottom, above, below, vertical, horizontal, clockwise, counterclockwise,upstream, downstream, forward, aft, etc.) are only used foridentification purposes to aid the reader's understanding of the presentinvention, and do not create limitations, particularly as to theposition, orientation, or use of the invention. Connection references(e.g., attached, coupled, connected, and joined) are to be construedbroadly and can include intermediate members between a collection ofelements and relative movement between elements unless otherwiseindicated. As such, connection references do not necessarily infer thattwo elements are directly connected and in fixed relation to oneanother. The exemplary drawings are for purposes of illustration onlyand the dimensions, positions, order and relative sizes reflected in thedrawings attached hereto can vary.

FIG. 1 is a schematic cross-sectional diagram of a gas turbine engine 10for an aircraft. The engine 10 has a generally longitudinally extendingaxis or centerline 12 extending forward 14 to aft 16. The engine 10includes, in downstream serial flow relationship, a fan section 18including a fan 20, a compressor section 22 including a booster or lowpressure (LP) compressor 24 and a high pressure (HP) compressor 26, acombustion section 28 including a combustor 30, a turbine section 32including a HP turbine 34, and a LP turbine 36, and an exhaust section38.

The fan section 18 includes a fan casing 40 surrounding the fan 20. Thefan 20 includes a plurality of fan blades 42 disposed radially about thecenterline 12. The HP compressor 26, the combustor 30, and the HPturbine 34 form a core 44 of the engine 10, which generates combustiongases. The core 44 is surrounded by core casing 46, which can be coupledwith the fan casing 40.

A HP shaft or spool 48 disposed coaxially about the centerline 12 of theengine 10 drivingly connects the HP turbine 34 to the HP compressor 26.A LP shaft or spool 50, which is disposed coaxially about the centerline12 of the engine 10 within the larger diameter annular HP spool 48,drivingly connects the LP turbine 36 to the LP compressor 24 and fan 20.The spools 48, 50 are rotatable about the engine centerline and coupleto a plurality of rotatable elements, which can collectively define arotor 51.

The LP compressor 24 and the HP compressor 26 respectively include aplurality of compressor stages 52, 54, in which a set of compressorblades 56, 58 rotate relative to a corresponding set of staticcompressor vanes 60, 62 (also called a nozzle) to compress or pressurizethe stream of fluid passing through the stage. In a single compressorstage 52, 54, multiple compressor blades 56, 58 can be provided in aring and can extend radially outwardly relative to the centerline 12,from a blade platform to a blade tip, while the corresponding staticcompressor vanes 60, 62 are positioned upstream of and adjacent to therotating blades 56, 58. It is noted that the number of blades, vanes,and compressor stages shown in FIG. 1 were selected for illustrativepurposes only, and that other numbers are possible.

The blades 56, 58 for a stage of the compressor can be mounted to a disk61, which is mounted to the corresponding one of the HP and LP spools48, 50, with each stage having its own disk 61. The vanes 60, 62 for astage of the compressor can be mounted to the core casing 46 in acircumferential arrangement.

The HP turbine 34 and the LP turbine 36 respectively include a pluralityof turbine stages 64, 66, in which a set of turbine blades 68, 70 arerotated relative to a corresponding set of static turbine vanes 72, 74(also called a nozzle) to extract energy from the stream of fluidpassing through the stage. In a single turbine stage 64, 66, multipleturbine blades 68, 70 can be provided in a ring and can extend radiallyoutwardly relative to the centerline 12 while the corresponding staticturbine vanes 72, 74 are positioned upstream of and adjacent to therotating blades 68, 70. It is noted that the number of blades, vanes,and turbine stages shown in FIG. 1 were selected for illustrativepurposes only, and that other numbers are possible.

The blades 68, 70 for a stage of the turbine can be mounted to a disk71, which is mounted to the corresponding one of the HP and LP spools48, 50, with each stage having a dedicated disk 71. The vanes 72, 74 fora stage of the compressor can be mounted to the core casing 46 in acircumferential arrangement.

Complementary to the rotor portion, the stationary portions of theengine 10, such as the static vanes 60, 62, 72, 74 among the compressorand turbine section 22, 32 are also referred to individually orcollectively as a stator 63. As such, the stator 63 can refer to thecombination of non-rotating elements throughout the engine 10.

In operation, the airflow exiting the fan section 18 is split such thata portion of the airflow is channeled into the LP compressor 24, whichthen supplies pressurized air 76 to the HP compressor 26, which furtherpressurizes the air. The pressurized air 76 from the HP compressor 26 ismixed with fuel in the combustor 30 and ignited, thereby generatingcombustion gases. Some work is extracted from these gases by the HPturbine 34, which drives the HP compressor 26. The combustion gases aredischarged into the LP turbine 36, which extracts additional work todrive the LP compressor 24 and fan 18, and the exhaust gas is ultimatelydischarged from the engine 10 via the exhaust section 38. The driving ofthe LP turbine 36 drives the LP spool 50 to rotate the fan 20 and the LPcompressor 24.

A portion of the pressurized airflow 76 can be drawn from the compressorsection 22 as bleed air 77. The bleed air 77 can be drawn from thepressurized airflow 76 and provided to engine components requiringcooling. The temperature of pressurized airflow 76 entering thecombustor 30 is significantly increased. As such, cooling provided bythe bleed air 77 is necessary for operating of such engine components inthe heightened temperature environments.

A remaining portion of the airflow 78 bypasses the LP compressor 24 andengine core 44 and exits the engine assembly 10 through a stationaryvane row, and more particularly an outlet guide vane assembly 80,comprising a plurality of airfoil guide vanes 82, at the fan exhaustside 84. More specifically, a circumferential row of radially extendingairfoil guide vanes 82 are utilized adjacent the fan section 18 to exertsome directional control of the airflow 78.

A side view of a turbine assembly 99 of the turbine engine isillustrated in FIG. 2. The turbine assembly 99 comprises acircumferential band, such as an inner band 106 or outer band 108,wherein the inner band 106 is radially spaced inwardly from the outerband 108 and a flow path can be defined between the bands 106, 108. Thebands 106, 108 can each have an upstream edge 110, a downstream edge112, and a surface 114 extending between the upstream and downstreamedges 110, 112. A component, illustrated herein as an airfoil 104 (forexample, a blade 68 or vane 72) can extend from at least one of theinner band 106 and outer band 108 into the flow path, and can also havea leading edge that confronts the flow path. In addition, a spacingdistance 116 can be constructed between each blade 68 and adjacent vane72. It should be understood that this blade-vane pair was selected forillustrative purposes only and is not meant to be limiting.

It is further contemplated that the component may comprise a temperatureprobe, strut, nozzle, pyrometer, other instrumentation probe, or anyother device that can extend from either or both of the bands 106, 108and have a leading edge confronting the flow path.

FIG. 3 shows a close up view of the interface between the vane 72 andblade 68. A gap 138 can exist in this interface, and other components ormaterials (not shown) may be positioned within or below the gap 138.

FIG. 4 shows a perspective view of a simplified lower half of a nozzleairfoil 104 according to a first embodiment of the invention. Theairfoil 104 comprises a pressure side 122 and a suction side 124. Theairfoil 104 can extend from a leading edge 126 to a trailing edge 128 todefine a chord-wise direction, and also extend from a root 130 to a tip132 to define a span-wise direction where the root 130 is proximate thesurface 114 of at least one of the bands 106, 108. The leading edge 126can have an effective radius 134 and be positioned an axial distance 136away from the upstream edge 110 of the inner band 106 wherein thedistance 136 can be a function of the radius 134 of the leading edge126. The distance 136 can be approximately between 1.0 and 5.0 times theleading edge radius 134. In addition, a trough 200 can be located in thesurface 114, and can have an inner wall 202 at least along the leadingedge 126 of the airfoil 104. The trough 200 can also extend around theleading edge 126 to a least a portion of the suction side 124 and thepressure side 122. Further, at least one flow enhancer 300 can bepositioned within the trough 200 and can comprise at least one of aturbulator, fastback turbulator, pin fins, pin bank, vortex generator,or chevron in non-limiting examples.

A side sectional view of a portion of the airfoil 104 taken along line5-5 is illustrated in FIG. 5. At the root 130 of the leading edge 126 ofthe airfoil 104, the trough 200 can have a depth 204 and a width 206.Either or both of the depth 204 and width 206 can be a function of theradius 134 of the leading edge 126. For example, the depth 204 or width206 can be at least 0.5 times the radius 134 or up to 3.0 times theradius 134, and it should be understood that this example is given forillustrative purposes and is not meant to be limiting. Further, whilethe trough 200 is illustrated with a constant depth 204 and width 206,it is contemplated that the trough 200 may also be contoured withtapering width 206 or depth 206 as it extends around the pressure sideor suction side. In a preferred embodiment the inner wall 202 of thetrough 200 can be in vertical alignment with the leading edge 126, orthe inner wall could be positioned an approximate distance of 1.0 timesthe radius 134 upstream of the leading edge 126. Further, a highpressure combustion gas flow known as a bow wave 250 can be formed at anextent 252 upstream of the leading edge 126, and the bow wave 250 cantypically extend vertically upward as shown by the straight dashed line.

Turning to FIG. 6, a top view of the airfoil 104 is illustrated. Thetrough 200 can extend around the leading edge 126 of the airfoil 104 toat least a portion of the suction side 124 and pressure side 122.

The dimensions of the trough 200 can be chosen such that air flowingtoward the leading edge 126 can move into the trough 200 and form avortex extending about the leading edge 126 from the pressure side 122to the suction side 124 as shown in FIG. 4. Exemplary air flow lines Fare shown in FIGS. 4 and 5 for air moving near the surface 114 andforming a vortex in the trough 200. This vortex can extend up theairfoil 104 along at least a portion of the span from the root 130, andcan also extend around the airfoil 104 along at least a portion of thepressure side 122 and the suction side 124. Moreover, the formation ofthe vortex can limit the extent 252 of the bow wave 250 at the root 130to be smaller than or equal to the width 206 of the trough 200 as shownby the curved dashed line in FIGS. 5 and 6, and can also be smaller thanor equal to the width 206 of the trough 200 at up to 15% of the spanfrom the root 130 as shown in FIG. 5.

In FIGS. 7 and 8, a portion of one of the airfoils 104 in the turbineassembly 99 from FIG. 2 is illustrated according to a second embodimentof the invention where the dimensions as described for the firstembodiment also apply to the second embodiment. The trough 200 cancontain a cove 210 located at least at the position of the leading edge126 of the airfoil 104 as seen in FIG. 7. Exemplary air flows F areshown for air moving toward the leading edge 126 of the airfoil 104 andinto the trough 200. The inner wall 202 of the trough 200 can berecessed by an amount 212 into the leading edge 126, forming the cove210 as seen in FIG. 8. The cove 210 is tantamount to an undercut, nook,or cavity within the trough 200 at the leading edge 126. Further, thecove 210 may extend around the pressure side 122 and suction side 124with a constant depth 204 and width 206, or it can also be contouredwith varying depth 204 and width 206.

A method of reducing the upstream extent 252 of a bow wave 250 from anairfoil 104 comprises forming a vortex at a leading edge 126 of theairfoil 104 along at least a portion of the root 130 of the airfoil 104.The vortex can be of sufficient strength to limit the upstream extent252 of the bow wave 250 to a distance less than or equal to the width206 of the trough 200 for at least 15% of the span from the root 130 asshown in FIG. 5.

The introduction of the trough 200 at the root 130 of the airfoil 104can act to suppress the broadcast of the bow wave 252; in one examplethe extent 252 of the bow wave 250 was reduced by 66% compared to acurrent design. Referring to FIG. 3, the gap 138 can be continuallypurged with cooling air to protect any temperature-sensitive materialspositioned within the cavity below the gap 138. The bow wave broadcastreduction can allow for the spacing distance 116 to be reduced betweeneach adjacent vane 72 and blade 68 while still preventing ingestion ofthe bow wave into the gap 138. It can be appreciated that a reduction inthe spacing distance 116 can allow for a reduction in the axial lengthof the turbine assembly 99, which can reduce the overall engine lengthand weight, manufacturing costs, external drag on the engine, and theamount of material to cool within the engine, and also increase theengine's efficiency.

It should be appreciated that application of the disclosed design is notlimited to turbine engines with fan and booster sections, but isapplicable to turbojets and turboshaft engines as well.

This written description uses examples to disclose the invention,including the best mode, and also to enable any person skilled in theart to practice the invention, including making and using any devices orsystems and performing any incorporated methods. The patentable scope ofthe invention is defined by the claims, and may include other examplesthat occur to those skilled in the art. Such other examples are intendedto be within the scope of the claims if they have structural elementsthat do not differ from the literal language of the claims, or if theyinclude equivalent structural elements with insubstantial differencesfrom the literal languages of the claims.

What is claimed is:
 1. An assembly fora turbine engine having acombustion section emanating a combustion airflow, comprising: radiallyspaced inner and outer bands defining a flow path between the inner andouter bands, wherein the inner and outer bands each have an upstreamedge, a downstream edge, and a surface extending between the upstreamedge and the downstream edge; an airfoil extending from one of the inneror outer bands into the flow path and defining a leading edgeconfronting the flow path, the airfoil having a pressure side and asuction side, and extending from the leading edge to a trailing edge todefine a chord-wise direction, and extending from a root to a tip todefine a span-wise direction, where the root is proximate one of theinner or outer bands and the tip is proximate the other of the inner orouter bands; and an imperforate trough extending into the surface of oneof the inner or outer bands along the leading edge; whereby theimperforate trough generates a vortex in the combustion airflow.
 2. Theassembly of claim 1 wherein the imperforate trough generates a vortexthat limits a bow wave generated by the leading edge to not extendupstream beyond the imperforate trough at up to 15% of the span from theroot.
 3. The assembly of claim 1 wherein the imperforate trough extendsaround the leading edge to at least a portion of the suction side andthe pressure side.
 4. The assembly of claim 1 wherein the airfoil is oneof a vane or blade in one of a compressor or turbine.
 5. The assembly ofclaim 1 wherein the imperforate trough generates a vortex that limits abow wave generated by the leading edge to not extend upstream beyond theimperforate trough at a portion of the airfoil proximate the surface. 6.The assembly of claim 1 wherein the imperforate trough has a depth thatis between 0.25 and 3.5 times a radius of the leading edge.
 7. Theassembly of claim 1 wherein the imperforate trough has a width that isbetween 0.25 and 3.5 times a radius of the leading edge.
 8. The assemblyof claim 1 wherein a distance from the upstream edge to leading edge isbetween 0.25 and 5.0 times a radius of the leading edge.
 9. The assemblyof claim 1 further comprising a flow enhancer located within theimperforate trough.
 10. The assembly of claim 9 wherein the flowenhancer comprises at least one of a turbulator, fastback turbulator,pin fins, pin bank, vortex generator, or chevron.
 11. The assembly ofclaim 1 further comprising a cove located at least along an intersectionof the leading edge and the imperforate trough.
 12. The assembly ofclaim 11 wherein the imperforate trough comprises an inner wall recessedinto the leading edge to form the cove.
 13. The assembly of claim 11wherein the cove comprises a cavity within the imperforate trough at theleading edge.
 14. The assembly of claim 11 wherein the cove extendsaround at least one of the pressure side or the suction side of theairfoil.
 15. An assembly for a turbine engine having a combustionsection emanating a combustion airflow, comprising: at least onecircumferential band at least partially defining an air flow path alongwhich the combustion airflow moves and having an upstream edge and adownstream edge relative to the air flow path, with a surface extendingbetween the upstream edge and the downstream edge; a component extendingfrom the at least one circumferential band into the airflow path anddefining a leading edge confronting the air flow path; an imperforatetrough extending into the surface of the at least one circumferentialband along the leading edge; and a flow enhancer located within theimperforate trough, wherein the flow enhancer comprises at least one ofa turbulator, fastback turbulator, pin fins, pin bank, vortex generator,or chevron; whereby the imperforate trough generates a vortex in thecombustion airflow.
 16. The assembly of claim 15 wherein the componentis one of a temperature probe, strut, nozzle, pyrometer, airfoil orother instrumentation probe.
 17. The assembly of claim 16 wherein thecomponent is an airfoil having a pressure side and a suction side, andextending from the leading edge to a trailing edge to define achord-wise direction, and extending from a root to a tip to define aspan-wise direction, with one of the root and the tip proximate the atleast one circumferential band.
 18. The assembly of claim 17 wherein theat least one circumferential band comprises radially spaced inner andouter bands defining a flow path between the inner and outer bands,wherein the inner bands and outer bands each have an upstream edge, adownstream edge, and a surface extending between the upstream edge andthe downstream edge, and the root is proximate one of the inner or outerbands and the tip is proximate the other of the inner or outer bands.19. An assembly for a turbine engine having a combustion sectionemanating a combustion airflow, comprising: at least one circumferentialband at least partially defining an air flow path along which thecombustion airflow moves and having an upstream edge and a downstreamedge relative to the air flow path, with a surface extending between theupstream edge and the downstream edge; a component extending from the atleast one circumferential band into the air flow path and defining aleading edge confronting the air flow path; and an imperforate troughextending into the surface of the at least one circumferential bandalong the leading edge; and a cove located at least along anintersection of the leading edge and the imperforate trough; whereby theimperforate trough generates a vortex in the combustion airflow.
 20. Theassembly of claim 19 wherein the imperforate trough comprises an innerwall recessed into the leading edge to form the cove.
 21. The assemblyof claim 19 wherein the cove comprises a cavity within the imperforatetrough at the leading edge.
 22. The assembly of claim 19 wherein thecove extends around at least one of a pressure side or a suction side ofthe component.
 23. The assembly of claim 19 further comprising a flowenhancer located within the imperforate trough or the cove.
 24. Theassembly of claim 23 wherein the flow enhancer comprises at least one ofa turbulator, fastback turbulator, pin fins, pin bank, vortex generator,or chevron.